In conventional axial compressors of turbomachines, such as aircraft engines and static gas turbines, a degree of reaction in the front compressor area drops from an inlet-side maximum to a minimum. The degree of reaction may be calculated in a simplified way according to the following formula, based on metal angles:
  R  =            tan      ⁢                          ⁢                                    β            ⁢                                                  ⁢            1                    +          β2                2                            tan        ⁢                                  ⁢                              β1            +            β2                    2                    +              tan        ⁢                                                                    ⁢                          α1              +              α2                                2                    
The angles α1, α2, β1, β2 are, as shown in FIG. 1, marked between tangents of the respective camber line and an axial flow direction x of the turbomachine. α1 is marked from the trailing edge of a guide blade 2 of a row of guide blades n−1. α2 is marked toward the leading edge of a guide blade 4 of a row of guide blades n. β1 is marked toward the leading edge of a moving blade 6 of a row of moving blades n. β2 is marked from the trailing edge of moving blade 6 of a row of moving blades n. Moving blade 6 or row of moving blades n thereby passes through between guide blades 2, 4 or rows of guide blades n−1, n in circumferential direction u. Letter n designates whole number multiples of 1, 2, etc. Usually, the degree of reaction for compressors covers a range between 0.5 and 1.0. Turbines usually cover degrees of reaction from 0.0 . . . 0.05 through 0.5. The drop of the degree of reaction upstream from the first not adjustable guide baffle is followed by a rise of the degree of reaction up to the compressor outlet in known turbomachines. Due to the rise, a residual swirl in the main flow may be reduced in the rear compressor area and in particular at the outlet guide baffle of the compressor; however, the rise causes a high load on the rear stages, which results in stability and efficiency limitations. In order to achieve a required stability, the rear guide baffles require a high number of blades.
It is the object of the present invention to create a method for operating a compressor of a turbomachine, with the aid of which a high stability and a high efficiency may be achieved. In addition, it is the object of the present invention to create a compressor of a turbomachine which has both a high stability and a high efficiency, and to create a turbomachine with a high stability and a high efficiency.